Aircraft control system and method

ABSTRACT

An aircraft control system includes a co-axial, counter-rotating propeller shaft assembly. Also included is a first rotor operatively coupled to the propeller shaft assembly, the first rotor having a first plurality of blades mounted thereto, wherein the first plurality of blades are disposed at a substantially identical nominal pitch during rotation of the first rotor. Further included is a second rotor operatively coupled to the propeller shaft assembly, the second rotor having a second plurality of blades mounted thereto, wherein a pitch of the second plurality of blades is configured to cyclically change during rotation of the second rotor.

BACKGROUND OF THE INVENTION

The embodiments herein relate to aircrafts and, more particularly, to anaircraft control system, as well as a method of controlling an aircraft.

Design of rotors and propellers is quite complex. A large number offactors must be taken into account, including flexure of the rotor underheavy loads and the required motions of the rotor blades with respect tothe drive mechanism.

Rigid turboprop propeller systems provide collective pitch control ofthe propeller blades. Pitch angles ranging from a fully featheredminimum drag angle to pitch angles which provide reverse thrust aretypically provided to provide propeller speed and power management.Inflow angles not along the axis of rotation due to aircraft maneuversgenerate bending moments on the propeller shaft and subsequent twistingof the airframe. The resulting bending moments can be rather large andconventional propeller systems are therefore rigidly structured.

Fully articulated rotors such as those of helicopters provide cyclic andcollective pitch of the rotor blades. Articulation of the rotor discplane vectors the rotor thrust to provide fore, aft and lateral movementof the helicopter with minimal bending moment of the rotor shaft. Ascompared to rigid turboprop propeller systems, articulated rotor systemsof a helicopter are significantly more complex. As such, the controlbenefits associated with fully articulated rotors are offset by thedesign and operational complexity.

BRIEF DESCRIPTION OF THE INVENTION

According to one embodiment, an aircraft control system includes aco-axial, counter-rotating propeller shaft assembly. Also included is afirst rotor operatively coupled to the propeller shaft assembly, thefirst rotor having a first plurality of blades mounted thereto, whereinthe first plurality of blades are disposed at a substantially identicalnominal pitch during rotation of the first rotor. Further included is asecond rotor operatively coupled to the propeller shaft assembly, thesecond rotor having a second plurality of blades mounted thereto,wherein a pitch of the second plurality of blades is configured tocyclically change during rotation of the second rotor.

According to another embodiment, a method of controlling an aircraft isprovided. The method includes rotating a first rotor operatively coupledto a propeller shaft assembly in a first direction. The method alsoincludes rotating a second rotor operatively coupled to the propellershaft assembly in a second direction that is opposite of the firstdirection. The method further includes maintaining a first plurality ofblades mounted to the first rotor at a substantially identical nominalpitch during rotation of the first rotor. The method yet furtherincludes cyclically changing the pitch of a second plurality of bladesmounted to the second rotor during rotation of the second rotor. Themethod also includes generating a moment upon cyclically changing thepitch of the second plurality of blades to control the aircraft.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter which is regarded as the invention is particularlypointed out and distinctly claimed in the claims at the conclusion ofthe specification. The foregoing and other features and advantages ofthe invention are apparent from the following detailed description takenin conjunction with the accompanying drawings in which:

FIG. 1 is a general perspective view of a turboprop assembly driven byan engine;

FIG. 2 is a schematic force diagram of a co-axial, counter-rotatingpropeller assembly;

FIG. 3 is a schematic force diagram of an aircraft utilizing thepropeller assembly according to an exemplary embodiment; and

FIG. 4 is a flow diagram illustrating a method of controlling anaircraft with the propeller assembly.

DETAILED DESCRIPTION OF THE INVENTION

Referring to FIG. 1, a general perspective view of a propeller system 20is illustrated. It should be understood that although a propeller systemtypical of a turboprop aircraft is illustrated in the disclosedembodiment, various rigid prop/rotor systems including tilt rotor andtilt wing systems will benefit from the embodiments described herein.

A gas turbine engine (illustrated schematically at 22) which rotates aturbine output shaft 24 at a high speed powers the propeller system 20.The turbine output shaft 24 drives a gear reduction gearbox (illustratedsomewhat schematically at 26) which decreases shaft rotation speed andincreases output torque. The gearbox 26 drives a propeller shaftassembly 28 which rotates a propeller hub 30 and a plurality ofpropeller blades 32 which extend therefrom. It should be understood thata conventional offset gearbox will also benefit from the presentinvention. Axis A is substantially perpendicular to a plane P which isdefined by the plurality of propeller blades 32. It should be understoodthat an in-line and an offset gearbox will benefit from the presentinvention.

Referring to FIG. 2, the propeller shaft assembly 28 and associatedcomponents are further illustrated. In particular, the propeller shaftassembly 28 shown is a co-axial, counter-rotating propeller shaftassembly and the propeller system 20 operates as an aircraft controlsystem, as described in detail below. The propeller system 20 includes afirst rotor 34 having a first hub 36 operatively coupled thereto.Extending from the first hub is a first plurality of blades 38. A secondrotor 40 having a second hub 42 operatively coupled thereto is included,with a second plurality of blades 44 extending from the second hub 42.As noted above, the first rotor 34 and the second rotor 40, andtherefore the first plurality of blades 38 and the second plurality ofblades 44, are configured to rotate in opposite directions about acommon axis A with both rotors producing thrust in the same direction.As will be described in detail below, the second plurality of blades 44is configured to be capable of cyclically changing pitch during rotationof the second rotor 40. In contrast, each of the first plurality ofblades 38 is disposed at a substantially identical nominal pitch duringrotation of the first rotor 34. The nominal pitch of first the pluralityof blades 38 may change based on power input and operating conditions,for example, but each of the blades of the first rotor 34 are at thesubstantially identical pitch. The “pitch” of the blades is defined asthe rotational position of the blade about an axis from the root of theblade to the tip of the blade.

The arrows in FIG. 2 indicate the direction of the airflow. In theillustrated embodiment, the engine 22 driving the propeller system 20 isupstream of both rotors and the first rotor 34 is upstream of the secondrotor 40. This embodiment is commonly referred to as a “pusher”configuration. In an alternative embodiment, the direction of theairflow is reversed, such that the second rotor 40 is upstream of thefirst rotor 34, which are both upstream of the engine 22. Thisembodiment is commonly referred to as a “tractor” or “puller”configuration. Most importantly, and irrespective of anupstream-downstream configuration, it is less complex to accomplishcyclic pitch on the rotor that is closest to the engine. However, it iscontemplated that cyclic pitch is actuated on the rotor located furthestfrom the engine 22, either in addition to actuation of cyclic pitch onthe other rotor or in combination therewith.

In operation, the propeller system 20 generates a once per revolution(1P) variation in blade load through cyclic pitch of the first pluralityof blades 38. While the axis of the thrust vector remains perpendicularto the plane of the first plurality of blades 38, the variation in bladeload creates a bending moment on the propeller shaft assembly 28 whichappears fixed in relation to the aircraft. There is also a relativelysmall in-plane force generated due to the difference in torque onopposing blades. Such 1P variations may occur during aircraftmaneuvering when inflow angles are not on the propeller axis ofrotation. Conventional blade mounting arrangements accommodate theseoff-axis forces by rigidly mounting the propeller blades to the hub toprevent flapping and rigidly mount the propeller shaft assembly 28 tothe gearbox 26 (FIG. 1). Off-axis forces are thereby transmitteddirectly from the propeller blades to the airframe. The embodimentsherein advantageously utilize this conventional mounting arrangement togenerate aircraft attitude control through generation of a moment aboutthe propeller shaft 28 assembly (FIG. 1). Various structures and methodsrelating to cyclic pitch may be employed to facilitate cyclicallychanging the pitch of the second plurality of blades 44. For example,commonly owned U.S. Pat. No. 6,981,844, describes such a structure andmethod, the disclosure of which is incorporated by reference herein. Itis to be appreciated that the specific structure and method disclosed inthe above-referenced disclosure is not limiting of alternative cyclicpitch embodiments. It is to be understood that the general concept ofcyclically changing the pitch of propeller blades is applied to only onepropeller blade set of a co-axial, counter-rotating propeller shaftassembly. In particular, the pitch of the first plurality of blades 38is cyclically changed during rotation.

Referring to FIG. 3, an aircraft (illustrated schematically at 80) witha wing 82 for providing lift and one or more propeller systems 20 isshown according to one embodiment. The propeller systems 20 produceforward thrust and incorporate a pitch change actuator assembly 50. Asindicated, the propeller systems 20 provide Thrust (T1 & T2), andMoments (M1 & M2). Moments M1 & M2 are vectorally represented using theconventional “right hand rule” notation and may be directed anywhere,independently of each other 360 degrees within the plane of rotation ofthe second plurality of propeller blades.

The appropriate combination of the vectors M1 & M2 & T1 and T2 willproduce desired, roll, pitch and yaw moments Mx, My & Mz as desired tocontrol the pitch, roll and yaw of the aircraft 80. In addition, thethrust vectors T1 and T2 may be combined to contribute to the moment Mzon the aircraft to control the yaw as required. The roll is controlledby the coordinated application of a difference in the thrusts, T1 andT2, in combination with moments in the yaw direction. The resultantin-plane shear forces cause the aircraft to roll. Each of the momentsand vectors described above are provided by the incorporation ofdirectional cyclic pitch through the pitch change actuator assembly 50of the embodiments described herein in combination with the normalpropeller function of producing thrust for forward flight.

In the illustrated embodiment of FIG. 3, the propeller system 20 isillustrated proximate a wing of the aircraft 80, however, in analternative embodiment the propeller system 20 described in detail aboveis disposed proximate a nose of the aircraft 80. In such an embodiment,the co-axial, counter-rotating propeller shaft assembly 28 facilitatesvertical takeoff of the aircraft 80 from a “tail-sitter” position.Specifically, the aircraft 80 may be initially positioned in asubstantially vertical position and the control capabilities provided bythe propeller system 20 facilitates takeoff thrust and initialmaneuvering control of the aircraft 80 during takeoff from such aposition.

A method of controlling an aircraft 100 is also provided, as illustratedin FIG. 4 and with reference to FIGS. 1-3. The propeller system 20 and,more particularly, the co-axial, counter-rotating propeller shaftassembly 28, have been previously described and specific structuralcomponents need not be described in further detail. The method ofcontrolling an aircraft 100 includes rotating 102 the first rotor 34 ina first direction and rotating 104 the second rotor 40 in a seconddirection that is opposite of the first direction. Each of the firstplurality of blades 38 are maintained 106 at a substantially identicalnominal pitch during rotation of the first rotor 34. In contrast, thepitch of each of the second plurality of blades 44 is cyclically changed108 during rotation of the second rotor 40. As a result of cyclicallychanging the pitch of the second plurality of blades 44, a moment isgenerated 110 to control the aircraft 80.

While the invention has been described in detail in connection with onlya limited number of embodiments, it should be readily understood thatthe invention is not limited to such disclosed embodiments. Rather, theinvention can be modified to incorporate any number of variations,alterations, substitutions or equivalent arrangements not heretoforedescribed, but which are commensurate with the spirit and scope of theinvention. Additionally, while various embodiments of the invention havebeen described, it is to be understood that aspects of the invention mayinclude only some of the described embodiments. Accordingly, theinvention is not to be seen as limited by the foregoing description, butis only limited by the scope of the appended claims.

1. An aircraft control system comprising: a co-axial, counter-rotatingpropeller shaft assembly; a first rotor operatively coupled to thepropeller shaft assembly, the first rotor having a first plurality ofblades mounted thereto, wherein the first plurality of blades aredisposed at a substantially identical nominal pitch during rotation ofthe first rotor; and a second rotor operatively coupled to the propellershaft assembly, the second rotor having a second plurality of bladesmounted thereto, wherein a pitch of the second plurality of blades isconfigured to cyclically change during rotation of the second rotor. 2.The aircraft control system of claim 1, wherein the second rotor isdisposed downstream of the first rotor.
 3. The aircraft control systemof claim 1, wherein the second rotor is disposed upstream of the firstrotor.
 4. The aircraft control system of claim 1, wherein cyclicallychanging the pitch of the second plurality of blades generates a momentconfigured to control an aircraft.
 5. The aircraft control system ofclaim 1, wherein the first rotor and the second rotor are disposedproximate a fixed wing of an aircraft.
 6. The aircraft control system ofclaim 1, wherein the first rotor and the second rotor are disposedproximate a nose of an aircraft.
 7. The aircraft control system of claim6, wherein the aircraft is configured to take-off from a verticalorientation.
 8. A method of controlling an aircraft comprising: rotatinga first rotor operatively coupled to a propeller shaft assembly in afirst direction; rotating a second rotor operatively coupled to thepropeller shaft assembly in a second direction that is opposite of thefirst direction; maintaining a first plurality of blades mounted to thefirst rotor at a substantially identical nominal pitch during rotationof the first rotor; cyclically changing the pitch of a second pluralityof blades mounted to the second rotor during rotation of the secondrotor; and generating a moment upon cyclically changing the pitch of thesecond plurality of blades to control the aircraft.
 9. The method ofclaim 8, wherein the first rotor and the second rotor are rotatedproximate a nose of the aircraft about a common axis extending parallelto a longitudinal axis of the aircraft.
 10. The method of claim 8,wherein the first rotor and the second rotor are rotated proximate afixed wing of the aircraft about a common axis extending parallel to alongitudinal axis of the aircraft.